The present invention relates to the field of guidance, navigation, and control avionics, and, more particularly, to devices, systems, and methods for controlling the inertial attitude of an artificial, spinning satellite so as to support an accurate orbital navigation function and to support greater terrestrial coverage using the satellite.
Such small, artificial, spinning satellite constellations can collect data using, for example, optical or other sensor instrumentation. Necessarily, these data must be referenced to an inertial attitude sensing system that is an integrated part of the spinning satellite. Thus, attitude determination for both spinning and three-axis stabilized satellites is a critical operational function.
Historically, measurement of inertial attitude of an artificial satellite, a strategic missile, and the like has used an optical sensor, such as a star tracker, in combination with an accurate, inertial reference sensor suite consisting of at least three gyroscopes. See, for example, U.S. Pat. No. 6,577,929 to Johnson, et al., which is incorporated in its entirety herein by reference.
According to Johnson, et al., for higher accuracy and reliability, one classic form of an accurate gyroscope was a single-axis, floated, integrating-rate gyroscope. Briefly, inside each gyroscope, there is a rapidly-spinning wheel that is perpendicular inside a first can, which is floated inside a second can. An angular rate sensor disposed on the symmetric axis between the two cans is adapted to measure the angular precession rate due to one or more torques acting normal to the angular momentum vector along the spin motor axis effectively. Hence, an inertial reference sensor suite of three single-axis gyroscopes has been used to provide a complete inertial attitude reference, i.e., roll, pitch and yaw.
Typically, the gyroscopes are structured and arranged generally orthogonally to one another, to measure roll, pitch, and yaw (rates and angles) with a certain amount of long-term angle drift error. Periodic inertial attitude updates by the star tracker can bound the effect of the gyroscopes' drift errors. Conventional artificial satellites, strategic missiles, and the like are nominally not spinning, but may have a random attitude drift. Moreover, the inertial reference sensor is nominally not strapped to the artificial satellite, strategic missile, and the like, but is gimbaled so that the gyroscope sensors can be maintained in “inertial space” for better performance.
For some artificial satellites such as NASA's Apollo, the star tracker is gimbaled to the artificial satellite separately from the gyroscope inertial package. For some artificial satellites, the gyroscope and/or the star tracker may be strapped to the “low-attitude rate” satellite frame. While, for some strategic missiles, such as the surface-launched ballistic missile (SLBM), the star tracker can be mounted on a gyroscope-stabilized platform, oriented at stars or other astronomical objects through a “window” in the gimbals.
None of the above-mentioned concepts, however, is considered to be very small, light-weight, and/or ultra-low-power. Consequently, none of the above-mentioned concepts is considered to be useable on a class of small, light-weight, ultra-low-power spinning satellites.
Furthermore, critical issues that must be included in or accounted for in design include radiation susceptibility, temperature susceptibility, e.g., susceptibility to extreme temperature magnitudes and temperature gradients, and dynamic motion susceptibility. For example, to achieve a reliable attitude determination system suitable for a long-duration space vehicle that will operate over the South Atlantic Anomaly and/or the Van Allen Belts, optical sensors that are susceptible to total dose and/or single event latch-up radiation effects, e.g., a CCD or an APS sensor, are undesirable.
Accordingly, elimination of MEMS instrument gyroscopes, which include electronic components that are not radiation-hard, especially in a relatively high-radiation, relatively high-dynamic temperature, relatively high-dynamic acceleration environment, is desirable. Additional reasons for replacing MEMS gyroscopes include the extensive, real-time, calibration and compensation requirement associated with MEMS gyroscopes and their inherently high drift rates.
Therefore, it would be desirable to provide methods and systems for accurately determining inertial attitude of an artificial spinning satellite, a strategic spinning missile, and the like and, additionally, to provide methods and systems for controlling or adjusting the inertial attitude of the artificial spinning satellite, strategic spinning missile, and the like. Moreover, it would be desirable to provide such methods and systems that also reduce volume, weight, and power requirements.